In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot combustion gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the airfoils of the turbine vanes and turbine blades, upon which the hot combustion gases impinge. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to about 1800-2100° F.
Many approaches have been used to increase the operating temperature limit of the turbine blades and vanes to their current levels. For example, the composition and processing of the base materials themselves have been improved. Solidification is controlled to produce directionally oriented structures that orient the grain boundaries parallel to the major stress axis or eliminate the grain boundaries entirely, and also take advantage of the most suitable crystallographic orientations of the metal. Physical cooling techniques are used, such as providing internal cooling passages through which cooling air is passed.
In another approach, protective coatings are applied to the internal and external surfaces of the airfoils of the turbine blades and vanes. Aluminide diffusion coatings are used for the internal surfaces, and aluminide diffusion coatings or overlay coatings are used on the external surfaces. A ceramic thermal barrier coating layer may also overlie the aluminum-containing coating on the external surfaces.
Although these internal and external protective layers provide improved resistance to environmental damage of the turbine components and the ability to operate at higher temperatures, there is an opportunity for improvement. Thus, there is a need for improved protective coating systems and methods for their application that extend the capabilities of the turbine components even further. The present invention fulfills this need, and further provides related advantages.